Methods and apparatus for cooling gas turbine nozzles

ABSTRACT

A method for assembling a turbine nozzle for a gas turbine engine facilitates improving cooling efficiency of the turbine nozzle. The method includes providing a hollow doublet including a leading airfoil and a trailing airfoil coupled by at least one platform, wherein each airfoil includes a first sidewall and a second sidewall that extend between a respective leading and trailing edge. The method also includes inserting an insert into at least one of the airfoils, wherein the insert includes a first sidewall including a first plurality of cooling openings that extending therethrough, and a second sidewall including a second plurality of cooling openings extending therethrough, and wherein the first plurality of cooling openings facilitate more cooling of the airfoil than the second plurality of cooling openings.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engine nozzles and moreparticularly, to methods and apparatus for cooling gas turbine enginenozzles.

Gas turbine engines include combustors which ignite fuel-air mixtureswhich are then channeled through a turbine nozzle assembly towards aturbine. At lease some known turbine nozzle assemblies include aplurality of nozzles arranged circumferentially and configured asdoublets. A turbine nozzle doublet includes a pair ofcircumferentially-spaced hollow airfoil vanes coupled byintegrally-formed inner and outer band platforms.

The doublet type turbine nozzles facilitate improving durability andreducing leakage in comparison to non-doublet turbine nozzles.Furthermore, turbine nozzle doublets also facilitate reducingmanufacturing and assembly costs. In addition, because such turbinenozzles are subjected to high temperatures and may be subjected to highmechanical loads, at least some known doublets include an identicalinsert installed within each airfoil vane cavity to distribute coolingair supplied internally to each airfoil vane. The inserts include aplurality of openings extending through each side of the insert.

In a turbine nozzle, the temperature of the external gas is higher onthe pressure-side than on the suction-side of each airfoil vane. Becausethe openings are arranged symmetrically between the opposite insertsides, the openings facilitate distributing the cooling air throughoutthe airfoil vane cavity to facilitate achieving approximately the sameoperating temperature on opposite sides of each airfoil. However,because of the construction of the doublet, mechanical loads and thermalstresses may still be induced unequally across the turbine nozzle. Inparticular, because of the orientation of the turbine nozzle withrespect to the flowpath, typically the mechanical and thermal stressesinduced to the trailing doublet airfoil vane are higher than thoseinduced to the leading doublet airfoil vane. Over time, continuedoperation with an unequal distribution of stresses within the nozzle mayshorten a useful life of the nozzle.

BRIEF SUMMARY OF THE INVENTION

In one aspect of the invention, a method for assembling a turbine nozzlefor a gas turbine engine is provided. The method includes providing ahollow doublet including a leading airfoil vane and a trailing airfoilvane coupled by at least one platform, wherein each airfoil vaneincludes a first sidewall and a second sidewall that extend between arespective leading and trailing edge. The method also includes insertingan insert into at least one of the airfoil vanes, wherein the insertincludes a first sidewall including a first plurality of coolingopenings that extend therethrough, and a second sidewall including asecond plurality of cooling openings extending therethrough.

In another aspect, a method of operating a gas turbine engine isprovided. The method includes directing fluid flow through the engineusing at least one turbine airfoil nozzle that includes a leadingairfoil and a trailing airfoil coupled by at least one platform that isformed integrally with the leading and trailing airfoils, and whereineach respective airfoil includes a first sidewall and a second sidewallthat extend between respective leading and trailing edges to define acavity therein. The method also includes directing cooling air into theturbine airfoil nozzle such that the nozzle trailing airfoil is cooledmore than the leading airfoil.

In a further aspect of the invention, a turbine nozzle for a gas turbineengine is provided. The nozzle includes a pair of identical airfoilvanes coupled by at least one platform formed integrally with theairfoil vanes. Each airfoil vane includes a first sidewall and a secondsidewall that are connected at a leading edge and a trailing edge, suchthat a cavity is defined therebetween. The nozzle also includes at leastone insert that is configured to be inserted within the airfoil vanecavity and includes a first sidewall and a second sidewall. The insertfirst sidewall includes a first plurality of openings extendingtherethrough for directing cooling air towards at least one of theairfoil vane first and second sidewalls. The insert second sidewallincludes a second plurality of openings that extend therethrough fordirecting cooling air towards at least one of the airfoil vane first andsecond sidewalls. The first plurality of openings are configured tofacilitate lower metal temperatures therefrom than the second pluralityof openings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of a gas turbine engine;

FIG. 2 is an exploded perspective forward-looking-aft view of turbinenozzle that may be used with the gas turbine engine shown in FIG. 1; and

FIG. 3 is an exploded perspective aft-looking-forward view of theturbine nozzle shown in FIG. 2.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic illustration of a gas turbine engine 10 includinga fan assembly 12, a high-pressure compressor 14, and a combustor 16.Engine 10 also includes a high-pressure turbine 18 and a low-pressureturbine 20. Engine 10 has an intake, or upstream, side 28 and anexhaust, or downstream, side 30. In one embodiment, engine 10 is aCF6-80 engine commercially available from General Electric AircraftEngines, Cincinnati, Ohio.

In operation, air flows through fan assembly 12 and compressed air issupplied to high-pressure compressor 14. The highly compressed air isdelivered to combustor 16. Airflow from combustor 16 is dischargedthrough a turbine nozzle assembly (not shown in FIG. 1) that includes aplurality of nozzles (not shown in FIG. 1) and used to drive turbines 18and 20. Turbine 20, in turn, drives fan assembly 12, and turbine 18drives high-pressure compressor 14.

FIG. 2 is an exploded perspective forward-looking-aft view of turbinenozzle 50 that may be used with gas turbine engine 10 (shown in FIG. 1).FIG. 3 is an exploded perspective aft-looking-forward view of turbinenozzle 50. Nozzle 50 is known as a doublet and includes a pair ofcircumferentially-spaced airfoil vanes 52 coupled together by an arcuateradially outer band or platform 56 and an arcuate radially inner band orplatform 54. More specifically, in the exemplary embodiment, each band54 and 56 is formed integrally with airfoil vanes 52.

Inner band 54 includes a retention flange 60 that extends radiallyinwardly therefrom. More specifically, flange 60 extends substantiallyperpendicularly from band 54 with respect to a radially outer surface 62of flange 60. Outer band 56 also includes a retention flange 64 thatextends radially outwardly therefrom, and a leading edge flange 66 thatalso extends radially outwardly therefrom. More specifically, outer bandretention flange 64 and leading edge flange 66 extend substantiallyperpendicularly from band 56 with respect to a radially inner surface 68of band 56. Surfaces 62 and 68 define a radially outer and radiallyinner boundary for a flowpath through nozzle 50.

Airfoil vanes 52 are identical and include a leading airfoil vane 76 anda trailing airfoil vane 78. Each airfoil vane 52 includes a firstsidewall 80 and a second sidewall 82. First sidewall 80 is convex anddefines a suction side of each airfoil vane 76 and 78, and secondsidewall 82 is concave and defines a pressure side of each airfoil vane76 and 78. Sidewalls 80 and 82 are joined at a leading edge 84 and at anaxially-spaced trailing edge 86 of each airfoil vane 76 and 78. Morespecifically, each airfoil trailing edge 86 is spaced chordwise anddownstream from each respective airfoil leading edge 84.

First and second sidewalls 80 and 82, respectively, extendlongitudinally, or radially outwardly, in span from radially inner band54 to radially outer band 56. Additionally, first and second sidewalls80 and 82, respectively, define a cooling chamber 90 within each airfoilvane 52. More specifically, chamber 90 is bounded by an inner surface 92and 94 of each respective sidewall 80 and 82, and extends through eachband 54 and 56.

Each cooling chamber 90 is sized to receive an insert 100 therein. Morespecifically, lead airfoil chamber 90 is sized to receive a lead insert102, and trailing airfoil chamber 90 is sized to receive a trailinginsert 104 therein. Inserts 102 and 104 are substantially similar andeach includes a respective key feature 110 and 112, and an identicalattachment flange 114. Flange 114 extends from a radially outer end 116of each insert 102 and 104, and enables each insert 102 and 104 to besecured within each respective cooling chamber 90. In one embodiment,flange 114 is brazed to radially outer band 56. In another embodiment,flange 114 is welded to radially outer band 56.

Key features 110 and 112 extend through flange 114 at each insertradially outer end 116. Specifically, key features 110 and 112 areunique to each respective insert 102 and 104, and are sized to bereceived in a mating slot (not shown) that extends through nozzleradially outer band 56. More specifically, key features 110 and 112prevent lead insert 102 from being inadvertently inserted withintrailing airfoil vane 78, and prevent trailing insert 104 from beinginadvertently inserted within leading airfoil vane 76.

Each insert 102 and 104 has a cross sectional profile that issubstantially similar to that of a respective airfoil vane 76 and 78.More specifically, each insert 102 and 104 includes a first sidewall 120and 122, respectively, and a second sidewall 124 and 126. Accordingly,each insert first sidewall 120 and 122 is adjacent each respectiveairfoil vane first sidewall 80 when each insert 102 and 104 is installedwithin each respective cooling chamber 90. Each insert first sidewall120 and 122 is convex and defines a suction side of each respectiveinsert 102 and 104, and each insert second sidewall is concave anddefines a pressure side of each respective insert 102 and 104.Respective pairs of insert sidewalls 120 and 124, and 122 and 126, arejoined at respective leading edges 128 and 130, and at respectivetrailing edges 132 and 134.

Lead insert first sidewall 120 defines a suction side of lead insert 102and includes a first plurality of openings 140 that extend therethroughto a cavity 142 defined therein. Lead insert second sidewall 124includes a second plurality of openings 144 that extend therethrough tocavity 142. First and second sidewall openings 140 and 144 of insert 102are biased to facilitate cooling a suction side 80 of lead airfoil vane76, more than a pressure side 82 of lead airfoil vane 76. In theexemplary embodiment, the plurality of first sidewall openings 140 aregreater than that required to achieve substantially equal surfacetemperatures when compared to the plurality of second sidewall openings144. The ratio of ninety first sidewall openings 140 to ninety-sevensecond sidewall openings 144 results in biased cooling and is incontrast to known inserts which have a ratio of seventy-six firstsidewall openings to one hundred thirty-seven second sidewall openingswhich results in cooling all four airfoil sidewalls substantiallyequally. In an alterative embodiment, the larger volume of air isfacilitated because insert first sidewall 120 includes openings 140which are larger in diameter than corresponding openings 144 extendingthrough insert second sidewall 124. It should be noted that thearrangement of openings 140 and 144 with respect to each respectivesidewall 120 and 124 is variable. Furthermore, the number and size ofopenings 140 and 144 is also variable.

Trailing insert first sidewall 122 defines a suction side of trailinginsert 104 and includes a first plurality of openings 150 that extendtherethrough to a cavity 152 defined therein. Trailing insert secondsidewall 126 includes a second plurality of openings 154 that extendtherethrough to cavity 152. First sidewall openings 150 permit a largervolume of cooling air to pass therethrough than second sidewall openings154. More specifically, insert 104 is biased to facilitate cooling asuction side 80 of trailing airfoil vane 78, more than a pressure side82 of trailing airfoil vane 78. In the exemplary embodiment, the largervolume of air is facilitated because the plurality of first sidewallopenings 150 outnumber the plurality of second sidewall openings 154.More specifically, in the exemplary embodiment, first sidewall 122includes one hundred forty-two openings 150, and second sidewall 126includes ninety-seven openings 154. In an alterative embodiment, thelarger volume of air is facilitated because insert first sidewall 122includes openings 150 which are larger in diameter than correspondingopenings 154 extending through insert second sidewall 126. It should benoted that the arrangement of openings 150 and 154 with respect to eachrespective sidewall 122 and 126 is variable. Furthermore, the number andsize of openings 150 and 154 is also variable.

Each nozzle 50 is in flow communication with a cooling system (notshown) that directs cooling air into each airfoil vane cooling chamber90 for internal cooling of nozzle airfoil vanes 52. Specifically, thecooling system directs cooling air into each airfoil vane insert 100,which in-turn, channels the cooling air for cooling airfoil vanes 52. Inaddition to being biased to facilitate cooling a suction side of eachrespective airfoil vane 76 and 78, nozzle inserts 100 are biased tofacilitate cooling trailing airfoil vane 78 more than lead airfoil vane76. More specifically, trailing insert openings 150 and 154 are biasedsuch that a larger volume cooling air is directed towards trailingairfoil vane 78 through trailing insert 104 than is directed throughlead insert 102 towards lead airfoil vane 76. In the exemplaryembodiment, the larger volume of air is facilitated because theplurality of trailing airfoil vane first sidewall openings 150 outnumberthe plurality of, lead airfoil vane first sidewall openings 140. In analternative embodiment, the larger volume of air is facilitated byvarying the size of trailing airfoil vane openings 150 in comparison tolead airfoil vane openings 140.

During operation, cooling air is routed through the cooling system intonozzle 50, which may not be thermally loaded or mechanically stressedequally between adjacent airfoil vanes 76 and 78. More specifically, dueto gas loading, thermal variations, and mechanical loading, moremechanical and thermal stresses are induced and transmitted throughtrailing airfoil vane 78 than through lead airfoil vane 76. Becausenozzle inserts 102 and 104 provide nozzle 50 with a cooling scheme thatmay be customized to particular applications, cooling air supplied tonozzle 50 is allocated more to a suction side 80 of the airfoil vanes 52than to a pressure side 82 of the airfoil vanes 52. Accordingly, ascooling air is channeled into nozzle 50, inserts 102 and 104 directcooling air towards a respective nozzle airfoil vane 76 and 78. Thecooling air exits outwardly from each nozzle airfoil vane 52 through aplurality of airfoil trailing edge openings (not shown), and thermalstresses induced within each individual airfoil vane 76 and 78 arefacilitated to be reduced. Furthermore, by biasing the cooling airflowto cool trailing airfoil vane 78 more than lead airfoil vane 76, thermalstresses across nozzle 50 are facilitated to be controlled. As a result,although a maximum temperature on each airfoil vane concave surface isincreased, the thermal stresses induced in nozzle 50 are facilitated tobe controlled to counteract the mechanical stresses, thus facilitatingincreasing a useful life of nozzle 50.

The above-described turbine nozzle includes a pair of inserts thatenable a cooling scheme for the nozzle to be customized to particularapplications. Specifically, the inserts bias the distribution of coolingair supplied to the nozzle more to the suction side of each of theairfoil vanes, and more to the trailing airfoil vane in the doublet. Asa result, the inserts facilitate controlling thermal stresses inducedwithin the nozzle, and thus, facilitate increasing the useful life ofthe nozzle in a cost-effective and reliable manner.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A method for assembling a turbine nozzle for a gas turbine engine,said method comprising: providing a hollow doublet including a leadingairfoil vane and a trailing airfoil vane coupled by at least oneplatform, wherein each airfoil vane includes a first sidewall and asecond sidewall that extend between a respective leading and trailingedge; inserting an insert into at least one of the airfoil vanes,wherein the insert includes a first sidewall including a first pluralityof cooling openings that extending therethrough, and a second sidewallincluding a second plurality of cooling openings extending therethrough,and wherein the first plurality of cooling openings facilitie coolingthe airfoil more than the second plurality of cooling openings;inserting second insert into the remaining airfoil vane, wherein thefirst and second inserts non-identical.
 2. A method in accordance withclaim 1 wherein each airfoil vane includes a pressure side and a suctionside, inserting an insert into at least one of the airfoil vanes furthercomprises inserting an insert into at least one of the airfoil vanes tofacilitate biasing cooling airfoil towards the suction side of theairfoil vane.
 3. A method in accordance with claim 1 wherein the firstsidewall of each airfoil vane is convex, and the second sidewall of eachairfoil vane is concave, inserting an insert into at least one of theairfoil vanes further comprises inserting an insert into at least one ofthe airfoil vanes to facilitate biasing cooling airfoil towards theconvex side of the airfoil vane.
 4. A method in accordance with claim 1wherein inserting an insert into at least one of the airfoil vanesfurther comprises inserting a first insert into the leading airfoil vaneand a second insert into the trailing airfoil vane to facilitate coolingthe trailing airfoil vane more than the leading airfoil vane.
 5. Amethod in accordance with claim 1 wherein inserting an insert into atleast one of the airfoil vanes further comprises inserting a firstinsert into the leading airfoil vane and a second insert into thetrailing airfoil vane to facilitate reducing thermal stresses within theairfoil nozzle.
 6. A method of operating a gas turbine engine, saidmethod comprising: directing fluid flow through the engine using atleast one turbine airfoil nozzle that includes a leading airfoil and atrailing airfoil coupled by at least one platform that is formedintegrally with the leading and trailing airfoils, and wherein eachrespective airfoil includes a first sidewall and a second sidewall thatextend between respective leading and trailing edges to define a cavitytherein; and directing cooling air into the turbine airfoil nozzle suchthat the nozzle trailing airfoil is cooled more than the leadingairfoil.
 7. A method in accordance with claim 6 wherein directingcooling air into the turbine airfoil nozzle further comprises directingairflow into each respective airfoil cavity through an insert installedwithin the turbine nozzle to facilitate reducing thermal stresses withinthe turbine airfoil nozzle.
 8. A method in accordance with claim 6wherein directing cooling air into the turbine airfoil nozzle furthercomprises directing airfoil through at least one insert installed withinthe turbine nozzle that includes a first plurality of cooling openingsin flow communication with the airfoil first sidewall, and a secondplurality of cooling openings in flow communication with the airfoilsecond sidewall, wherein the first plurality of cooling openingsfacilitate cooling the airfoil more than the second plurality of coolingopenings.
 9. A method in accordance with claim 8 wherein the firstsidewall defines a suction side of the respective airfoil, and thesecond sidewall defines a pressure side of the respective airfoil,directing cooling air into the turbine airfoil nozzle further comprisesbiasing airflow entering the airfoil with the insert towards the suctionside of the airfoil.
 10. A method in accordance with claim 8 wherein thefirst sidewall is convex, and the second sidewall is concave, directingcooling air into the turbine airfoil nozzle further comprises biasingairflow entering the airfoil with the insert towards the convex side ofthe airfoil.
 11. A method in accordance with claim 6 wherein directingcooling air into the airfoil nozzle further comprises directing airflowinto each respective airfoil through a pair of non-identical insertsinstalled within the turbine nozzle, such that the trailing airfoil isbiased to receive more cooling air flow than the leading airfoil.
 12. Aturbine nozzle for a gas turbine engine, said nozzle comprising: a pairof identical airfoil vanes coupled by at least one platform that isformed integrally with said airfoil vanes, each said airfoil vanecomprising a first sidewall and a second sidewall connected at a leadingedge and a trailing edge to define a cavity therebetween, said airflowvane first sidewall defines an airfoil vane suction side, said airfoilvane second sidewall defines an airfoil vane pressure side; and at leastone inset configured to be inserted within said airfoil vane cavity andcomprising a first sidewall and a second sidewall, said insert firstsidewall is adjacent said airfoil vane first sidewall, said insert firstsidewall comprising a first plurality of openings extending therethroughfor directing cooling air towards at least one of said airfoil vanefirst and second sidewalls, said insert second sidewall comprising asecond plurality of openings extending therethrough for directingcooling air towards at least one of said airfoil vane first and secondsidewalls, said first plurality of openings configured to facilitatemore vane sidewall cooling than said second plurality of openings, saidfirst plurality of cooling openings is greater than said insert secondplurality of cooling openings.
 13. A nozzle in accordance with claim 12wherein said airfoil vane first sidewall defines an airfoil vane suctionside, said airfoil vane second sidewall defines an airfoil vane pressureside, said at least one insert further configured to be inserted withinat least one airflow cavity such that said insert first sidewall isadjacent said airfoil vane first sidewall.
 14. A nozzle in accordancewith claim 13 12 wherein said airfoil vane first sidewall is convex,said airfoil vane second sidewall is concave, said insert furtherconfigured to facilitate cooling said airfoil vane first sidewall morethan said airfoil vane second sidewall. 15.A nozzle in accordance withclaim 13 12wherein said at least one insert further configured to beinserted such that said insert first sidewall is in flow communicationand adjacent said airfoil vane first sidewall, said insert firstsidewall is convex, said insert second sidewall is concave.
 16. A nozzlein accordance with claim 13 12 wherein said pair of airfoil vanesfurther comprise a leading airfoil vane and a trailing airfoil vane,said at least one insert further comprises a first insert installedwithin said leading airfoil vane, and a non-identical second insertinstalled within said tailing airfoil vane, said inserts configured tofacilitate cooling said trailing airfoil vane more than said leadingairfoil vane.
 17. A nozzle in accordance with claim 13 12 wherein saidat least one insert further configured to facilitate reducing thermalstresses within said nozzle.
 18. A turbine nozzle for a gas turbineengine, said nozzle comprising: a leading airfoil; and a trailingairfoil; and at least one platform that is formed integrally with saidleading and trailing airfoils, and wherein each respective airfoilcomprises a first sidewall and a second sidewall that extend betweenrespective leading and trailing edges to define a cavity therein; and atleast one insert inserted within said airfoil cavity, said turbinenozzle coupled to a cooling system configured to direct cooling air intothe turbine airfoil nozzle such that a portion of said trailing airfoilis cooled more than other portions of said trailing airfoil, and suchthat said trailing airfoil first sidewall is cooled more than saidleading airfoil first sidewall.
 19. A turbine nozzle in accordance withclaim 18 wherein said turbine airfoil nozzle is further configured toreceive cooling air such that a portion of the leading airfoil is cooledmore than other portions of the leading airfoil.
 20. A turbine nozzle inaccordance with claim 18 wherein said turbine nozzle is furthercomprises: a first insert configured to be inserted within one of theairfoil vanes, wherein the first insert comprises a first sidewallincluding a first plurality of cooling openings extending therethrough,and a second sidewall including a second plurality of cooling openingsextending therethrough, and wherein the first plurality of coolingopenings facilitate cooling the airfoil more than the second pluralityof cooling openings, and wherein the first plurality of cooling openingsis greater than the second plurality of cooling openings; and a secondinsert configured to be inserted within the remaining airfoil vane. 21.A turbine nozzle in accordance with claim 20 wherein said first andsecond inserts are identical.
 22. A turbine nozzle in accordance withclaim 20 wherein said first and second inserts are non-identical.
 23. Aturbine nozzle in accordance with claim 20 wherein said first sidewalldefines a pressure side of the respective airfoil, and said secondsidewall defines a suction side of the respective airfoil, said insertconfigured to bias cooling airflow entering the airfoil towards thesuction side of the airfoil.
 24. A turbine nozzle in accordance withclaim 20 wherein said first sidewall is concave, and said secondsidewall is convex, said insert configured to bias cooling airfoilentering the airfoil towards the concave side of the airfoil.
 25. Aturbine nozzle in accordance with claim 18 wherein said nozzle comprisesa pair of non-identical inserts configured to bias the cooling airdirected to the trailing airfoil more than the cooling air flow directedto the leading airfoil.